Turbine vane assembly with ceramic matrix composite vanes

ABSTRACT

A turbine section adapted for use in a gas turbine engine includes a turbine case, a turbine wheel, and a turbine vane assembly. The turbine case is made from metallic materials. The turbine wheel is housed in the case. The turbine vane assembly is fixed to the case and configured to smooth and redirect air moving along a primary gas path of the turbine section.

FIELD OF THE DISCLOSURE

The present disclosure relates generally to vane assemblies for gasturbine engines, and more specifically to vanes that compriseceramic-containing materials.

BACKGROUND

Gas turbine engines are used to power aircraft, watercraft, powergenerators, and the like. Gas turbine engines typically include acompressor, a combustor, and a turbine. The compressor compresses airdrawn into the engine and delivers high pressure air to the combustor.In the combustor, fuel is mixed with the high pressure air and isignited. Products of the combustion reaction in the combustor aredirected into the turbine where work is extracted to drive thecompressor and, sometimes, an output shaft. Left-over products of thecombustion are exhausted out of the turbine and may provide thrust insome applications.

Products of the combustion reaction directed into the turbine flow overairfoils included in stationary vanes and rotating blades of theturbine. The interaction of combustion products with the airfoils heatsthe airfoils to temperatures that require the airfoils to be made fromhigh-temperature resistant materials and/or to be actively cooled bysupplying relatively cool air to the vanes and blades. To this end, someairfoils for vanes and blades are incorporating composite materialsadapted to withstand very high temperatures. Design and manufacture ofvanes and blades from composite materials presents challenges because ofthe geometry and strength required for the parts.

SUMMARY

The present disclosure may comprise one or more of the followingfeatures and combinations thereof.

A turbine section adapted for use in a gas turbine engine may include acase made from metallic materials, a turbine wheel, and a turbine vaneassembly. The case may be shaped to extend around a central referenceaxis. The turbine wheel may be housed in the case. The turbine vaneassembly may be fixed to the case and may be configured to smooth andredirect air moving along a primary gas path of the turbine sectionahead of interaction with the turbine wheel.

In some embodiments, the turbine wheel may include a disk, a pluralityof blades, and a rotatable seal. The disk may be mounted for rotationabout the central reference axis relative to the case. The plurality ofblades may be coupled to the disk for rotation with the disk. Therotatable seal element may be coupled to the disk for rotation with thedisk.

In some embodiments, the turbine vane assembly may include a pluralityof composite aero vanes made from ceramic matrix composite material, aplurality of structural vanes made from metallic materials, and a staticseal element. The plurality of composite vanes may be shaped to provideinner and outer end walls defining the primary gas path as well asairfoils that extend across the primary gas path. The plurality ofstructural vanes may be shaped to provide airfoils that extend acrossthe primary gas path. The static seal element may cooperate with therotatable seal element of the turbine wheel to provide a seal forresisting movement of gasses across the seal when the turbine section isin use within a gas turbine engine.

In some embodiments, the static seal element may be fixed to theplurality of structural vanes so as to be in turn coupled to the casewhile remaining free for relative movement in relation to the compositeaero vanes. Accordingly, the composite aero vanes may be substantiallyfree from carrying mechanical loads applied by pressure on the staticseal element to the case. In some embodiments, the seal provided by therotatable seal element and the static seal element may be arrangedradially inward of the primary gas path.

In some embodiments, the static seal element may include a seal paneland at least one seal land. The seal panel may divide axially adjacentcompartments within the turbine section. The at least one seal land maybe engaged by the rotatable seal element.

In some embodiments, the rotatable seal element may include a knifering. The knife ring may engage the land of the static seal element.

In some embodiments, the plurality of structural vanes may each includean inner end wall, an outer end wall, and an airfoil. The inner end wallmay face the primary gas path. The outer end wall may be spaced radiallyfrom the inner end wall and may face the primary gas path. The airfoilmay extend from the inner end wall to the outer end wall across theprimary gas path.

In some embodiments, the plurality of structural vanes may each includea seal mount. The seal mount may extend radially inwardly from the innerend wall away from the primary gas path. The static seal element may befixed to the seal mount.

In some embodiments, the plurality of structural vanes may each includea case mount. The case mount may extend radially outwardly from theouter end wall away from the primary gas path. The case mount may engagethe case to couple the structural vane and the static seal element tothe case.

In some embodiments, the plurality of composite aero vanes may be eachcoupled to the case by a metallic spar. The spar may extend through theairfoils of the composite aero vanes.

In some embodiments, the seal element may be fixed to the seal mount byat least one fastener. In some embodiments, each of the plurality ofcomposite aero vanes may be spaced apart from the static seal element.

According to an aspect of the present disclosure, a turbine vaneassembly adapted for use in a gas turbine engine may include a pluralityof composite aero vanes, a plurality of structural vanes, and a staticseal element. The plurality of composite aero vanes may be made fromceramic matrix composite material. The plurality of structural vanes maybe made from metallic materials. The static seal element may be fixed tothe plurality of structural vanes and may be spaced apart from thecomposite aero vanes.

In some embodiments, each of the plurality of composite aero vanes mayinclude an inner end wall, an outer end wall, and an airfoil. The innerend wall may extend partway around a central reference axis. The outerend wall may be spaced radially from the inner end wall to define aprimary gas path therebetween. The airfoil may extend from the inner endwall to the outer end wall across the primary gas path.

In some embodiments, each of the plurality of structural vanes mayinclude airfoils. The airfoils may extend across the primary gas path.In some embodiments, the static seal element may be arranged radiallyinward of the primary gas path.

In some embodiments, the plurality of structural vanes may each includea seal mount. The seal mount may extend radially inwardly from the innerend wall away from the primary gas path. The static seal element may befixed to the seal mount.

In some embodiments, the plurality of structural vanes may each includea case mount. The case mount may extend radially outwardly from theouter end wall away from the primary gas path. The case mount mayconfigured to engage a case so that the static seal element may be fixedto the case via the plurality of structural vanes. In some embodiments,the static seal element may be fixed to the seal mount by at least onefastener.

In some embodiments, the plurality of structural vanes may each includean inner end wall, an outer end wall, and an airfoil. The inner end wallmay face the primary gas path. The outer end wall may spaced radiallyfrom the inner end wall and may face the primary gas path. The airfoilmay extend from the inner end wall to the outer end wall across theprimary gas path.

In some embodiments, the airfoil of each of the plurality of structuralvanes may formed to include a cooling air passageway and a plurality offilm cooling holes. The plurality of cooling holes may interconnect thecooling air passageway with the primary gas path.

In some embodiments, the plurality of composite aero vanes may be eachmounted to a metallic spar. The spar may extend radially through theassociated airfoil.

These and other features of the present disclosure will become moreapparent from the following description of the illustrative embodiments.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cutaway of a gas turbine engine that includes a fan, acompressor, a combustor, and a turbine that includes a plurality ofturbine wheel assemblies and turbine vane assemblies, the turbine vaneassemblies being shown in further detail in FIGS. 2-6;

FIG. 2 is a front elevation view of one turbine vane assembly includedin the gas turbine engine of FIG. 1 showing that the turbine vaneassembly includes a plurality of composite aero vanes, a plurality ofstructural vanes made from metallic materials, and a static seal elementmounted radially inward of the ceramic and structural vanes;

FIG. 3 is a front elevation view of a portion of the turbine vaneassembly shown in FIG. 2 illustrating that the structural vanescomprising metallic materials are spaced between and separate from thecomposite aero vanes and have cooling holes for discharging cooling airacross the metallic materials of the structural vanes;

FIG. 4 is a cross-sectional view of a portion of the turbine section ofthe gas turbine engine of FIG. 1 showing that the turbine vane assemblyis mounted between a first stage turbine wheel and a second stageturbine wheel to redirect gas moving from the first stage turbine wheeltoward the second stage turbine wheel and showing that the static sealelement is fixed to the plurality of structural vanes for cooperationwith a rotatable seal element of the second stage turbine wheel;

FIG. 5 is a detail view of the turbine section of FIG. 4 showing thatstructural vanes are coupled directly with the static seal element totransfer axial loading through the structural vanes;

FIG. 6 is detail view of the turbine section similar to FIG. 5 showingthat the composite aero vanes are floating relative to the static sealelement and that the composite aero vanes are only required to transferaerodynamic loads acting on the composite aero vanes through a spar to acorresponding turbine case;

FIG. 7 is a cross-sectional view of the seal mount vane of FIG. 5showing that the structural vanes comprise metallic materials; and

FIG. 8 is a cross-sectional view of the self-supporting vane of FIG. 6showing that the composite aero vanes comprise ceramic matrix compositematerial.

DETAILED DESCRIPTION OF THE DRAWINGS

For the purposes of promoting an understanding of the principles of thedisclosure, reference will now be made to a number of illustrativeembodiments illustrated in the drawings and specific language will beused to describe the same.

A turbine section 18 according to the present disclosure is adapted foruse in a gas turbine engine 10 as suggested in FIGS. 1 and 2. The gasturbine engine 10 includes a fan 12, a compressor 14, a combustor 16,and a turbine 18. The fan 12 generates thrust for propelling anaircraft. The compressor 14 compresses and delivers air to the combustor16. The combustor 16 mixes fuel with the compressed air received fromthe compressor 14 and ignites the fuel. The hot, high-pressure gasesfrom the burning fuel are directed into the turbine 18 where the turbine18 extracts work from the gases to drive the compressor 14 and the fan12. In other embodiments, the gas turbine engine 10 may include a shaft,turboprop, or gearbox in place of the fan 12.

The turbine section 18 includes a turbine case 20, turbine wheels 22,24, and a turbine vane assembly 26 as shown in FIGS. 1-3. The turbinecase 20 is made from metallic materials and shaped to extend around acentral reference axis 11. The turbine case 20 surrounds the othercomponents of the turbine section 18. The turbine wheels 22, 24 aremounted for rotating about the axis 11 in the case 20. The turbine vaneassembly 26 is fixed to the case 20 and is configured to smooth andredirect air moving along a primary gas path 25 of the turbine section18 ahead of interaction with the turbine wheel 24 and downstream ofturbine wheel 22.

The turbine wheel 24 includes a disk 28, a plurality of blades 30, and arotating seal element 32 as shown in FIG. 4. The disk 28 is mounted forrotation about the central reference axis 11 relative to the case 20.The plurality of blades 30 are coupled to the disk for rotation with thedisk 28. The rotatable seal element 32 is coupled to the disk 28 forrotation with the disk 28.

The turbine vane assembly 26 includes a plurality of composite aerovanes 34, a plurality of structural vanes 36, and a static seal element38 as shown in FIGS. 2-6. Each of the composite aero vanes 34 comprisesceramic matrix material adapted to withstand high temperatures. However,the composite aero vanes 34 may have relatively low strength compared tothe structural vanes 36, which comprise metallic materials. Thestructural vanes 36 provide structural strength to the turbine vaneassembly 26 by receiving the mechanical loads, such as the axial momentapplied by pressure on the static seal element 38. The static sealelement 38 is fixed to the plurality of structural vanes 36 so as to bein turn coupled to the case 20 while remaining free for relativemovement in relation to the composite aero vanes 34. Accordingly, thecomposite aero vanes 34 are substantially free from carrying mechanicalloads applied by pressure on the static seal element 38 to the case 20.

The plurality of composite aero vanes 34 are made from integrally-formedceramic matrix composite material as noted above. Each composite aerovane 34 is shaped to provide inner and outer end walls 48, 46 definingthe primary gas path 25 and at least one airfoil 50 that extends acrossthe primary gas path 25.

The plurality of structural vanes 36 are made from metallic materials asnoted above. Each structural vane 36 is shaped to provide inner andouter endwalls 74, 72 and at least one airfoil 76 that extend across theprimary gas path 25.

The static seal element 38 cooperates with the rotatable seal element 32of the turbine wheel 24 to provide a compartment seal 35 for resistingmovement of gasses across the compartment seal 35 when the turbinesection 18 is in use within a gas turbine engine 10. The compartmentseal 35 formed by the static seal element 38 and the rotating sealelement 32 seals between axially adjacent compartments 43, 44 resultingin a first pressure in the compartment 43 on the first stage turbinewheel 22 side and a second pressure in the compartment 44 on the secondstage turbine wheel 24 side. The first pressure is greater than thesecond pressure resulting in a difference of pressure on either side 43,44 of the static seal element 38. The difference of pressure causes apressure force to act on a seal panel 39 of the static seal component38. The pressure force results in an axial moment in the turbine vaneassembly 26.

In the illustrative embodiment, the static seal element 38 is a singleintegrally formed ring. In other embodiments, the static seal element 38may include a plurality of static seal element segments to form the ringshape. The segmented static seal element 38 may also be provided withany appropriate seal apparatus to seal between each of the segments.

The static seal element 38 includes a seal panel 39, seal lands 40, 41,and a fastener 42 as shown in FIGS. 4 and 5. The seal panel 39 dividesaxially adjacent compartments 43, 44 within the turbine section 18. Theseal lands 40, 41 extend axially aft and away from the seal panel 39. Atleast one seal land 40, 41 is engaged by the rotatable seal element 32of the turbine wheel 24. In the illustrative embodiment, both seal lands40, 41 are engaged with the rotating seal element 32 to seal between theupstream turbine wheel 22 and the downstream turbine wheel 24. Thefastener 42 extends away from the seal panel 39 and engages a seal mount78 of the structural vane 36. The static seal element 38 is fixed to theseal mount 78 by at least one fastener 42. In the illustrativeembodiment, the fastener 42 is a bolt/nut combination. In otherembodiments, the fastener 42 may be another suitable fastener 42 such asa pin, rivet, or integrated manufacturing retention (casting, welding,etc.). In other embodiments, the fastener 42 may be some other suitablemechanical joint such as a cross-key, a transition fit, or radialbirdmouth with engaged rail.

The fastener 42 of the static seal element 38 is not fixed to thecomposite aero vane 34 in the illustrative embodiment. Each of theplurality of composite aero vanes 34 is spaced apart from the staticseal element 38. In the illustrative embodiment, the static seal element38 is spaced apart from the inner end wall 48 of the composite aero vane34 to leave a space 45 between the static seal element 38 and the innerend wall 48 of the composite aero vane 34. In other embodiments, a sealmay be arranged in the space 45 to seal between the static seal element38 and the composite aero vane 34 and still allow relative movement ofthe static seal element 38 relative to the composite aero vane 34.

Turning again to the plurality of composite aero vanes 34, each of theplurality of composite aero vanes 34 includes an outer end wall 46, aninner end wall 48, an airfoil 50, and a vane mount unit 52 as shown inFIG. 6. The inner end wall 48 is spaced radially inward of the outer endwall 46. The airfoil 50 extends between and interconnects the outer endwall 46 and the inner end wall 48. The airfoil 50 is shaped to redirectair moving along the primary gas path 25 of the turbine section 18 thatextends radially from the outer end wall 46 to the inner end wall 48.The airfoil 50 is also shaped to include a vane cavity 51 extendingradially through the airfoil 50 and opens at the inner and outer endwalls 46, 48. The outer end wall 46 defines a radially outer boundary ofthe primary gas path 25 and the inner end wall 48 defines a radiallyinner boundary of the primary gas path 25. The vane mount unit 52 mountsthe composite aero vanes 34 to the turbine case 20 without engaging thestatic seal component 38.

The airfoil 50 includes a radial outer end 54, a radial inner end 55,and a body 56 as shown in FIG. 6. The radial outer end 54 extendsradially-outwardly past the outer end wall 46 outside the primary gaspath 25 in the illustrative embodiment. The radial inner end 55 isspaced apart from the radial outer end 54 relative to the axis 11 andextends radially-inwardly past the inner end wall 48 outside the primarygas path 25. The radial inner end 55 of the airfoil 50 of the compositeaero vane 34 engages the vane mount unit 52. The body 56 extendsradially entirely between and interconnects the radial outer end 54 andthe radial inner end 55.

The vane mount unit 52 of the composite aero vanes 34 includes a carrier60, a spar 62, and a clamp nut 64 as shown in FIG. 6. The spar 62 ismade from metallic materials and the metallic spar 62 extends throughthe vane cavity 51 of the airfoils 50 of the composite aero vanes 34. Insome embodiments, the spar 62 may be hollow and include cooling holes totransmit cooling air to the composite aero vane 34 and/or into theinter-disk cavity between the turbine wheels 22, 24. The spar 62 isconfigured to receive aerodynamic loads from the airfoil 50 of thecomposite aero vane 34 during use of the turbine section 18 in the gasturbine engine 10. The carrier 60 is made from metallic materials and iscoupled to the spar 62. The carrier 60 engages the turbine case 20 tocarry aerodynamic loads from the spar 62 to the turbine case 20. Theclamp nut 64 is located radially inward of the inner end wall 48 of thecomposite aero vane 34 and mates with the spar 62 to clamp the compositeaero vane 34 blocking radial movement of the composite aero vane 34relative to the axis 11.

In the illustrative embodiment, the clamp nut radially retains thecomposite aero vane 34 relative to the spar 62. In other embodiments,other methods to radially retain the composite aero vane 34 may be used,such as a pin, other fastener, or integrated manufacturing retention(casting, welding, etc.). In some embodiments, the spar 62 may couple tothe static seal element 38 directly.

The carrier 60 of the vane mount unit 52 includes a body panel 66, aforward mount hanger 68, and an aft mount rail 70 as shown in FIG. 6.The forward hanger 68 extends radially outward from the carrier bodypanel 66 at a forward end of the body panel 66 and is engaged with aforward bracket 90 of the turbine case 20. The aft mount rail 70 extendsradially outward form the carrier body panel 66 at an aft end of thebody panel 66 and is engaged with an aft bracket 92 of the turbine case20. The spar 62 couples to the body panel 66 of the carrier 60 inbetween the forward and aft attachment features 68, 70.

The spar 62 of the vane mount unit 52 is shaped to engage the airfoil 50of the composite aero vane 34 at a location radially outward or radiallyinward of the primary gas path 25. The spar 62 engages the airfoil 50 ofthe composite aero vane 34 to transfer aerodynamic loads of the airfoil50 to the spar 62 so that the spar 62 may carry the aerodynamic loads tothe turbine case 20.

Each of the structural vanes 36 includes an outer end wall 72, an innerend wall 74, an airfoil 76, a case mount 77, and a seal mount 78 asshown in FIG. 5. The inner end wall 74 faces the primary gas path 25.The outer end wall 72 is spaced radially from the inner end wall 74 andfaces the primary gas path 25. The airfoil 76 extends from the inner endwall 74 to the outer end wall 72 across the primary gas path 25. Thecase mount 77 extends radially outwardly from the outer end wall 72 awayfrom the primary gas path 25. The case mount 77 engages the case 20 tocouple each structural vane 36 of the plurality of structural vanes 36and static seal element 38 to the case 20. The seal mount 78 extendsradially inwardly from the inner end wall 74 away from the primary gaspath 25 and the static seal element 38 is fixed to the seal mount 78.

The airfoil 76 of the structural vanes 36 are formed to include acooling air passageway 79 and a plurality of film cooling holes 80 asshown in FIGS. 2-5 and 7. The cooling air passageway 79 extends radiallythrough the airfoil 76 of the structural vane 36 to transmit cooling airinto the inter-disk cavity between the turbine wheels 22, 24. Theplurality of film cooling holes 80 interconnect the cooling airpassageway 79 with the primary gas path 25. In some embodiments, theairfoil 76 may include a metallic impingement tube to direct cooling airand/or locally increase cooling effectiveness.

In the illustrative embodiment, the airfoil 76 further includes atrailing edge cooling feature 83 as shown in FIG. 7. The trailing edgecooling feature 83 cools a trailing edge of the airfoil 76 of thestructural vanes 36.

The case mount 77 of the structural vanes 36 includes a forward hanger84 and an aft rail 86 as shown in FIG. 5. The forward hanger 84 extendsradially outward from the outer end wall 72 of the structural vane 36relative to the axis 11. The aft rail 86 is axially spaced apart fromthe forward hanger 84 and extends radially outward from the outer endwall 72 of the structural vane 36 relative to the axis 11. The forwardhanger 84 engages the case 20 at a location forward of the structuralvane 36 and the aft rail 86 engages the case 20 at a location aft of thestructural vane 36. The forward hanger 84 and the aft rail 86 engage thecase 20 to couple the structural vane 36 to the case 20 and transferaxial loads from the static seal element 38, axial and/orcircumferential loads from the structural vane 36, and aerodynamic loadsfrom structural vane 36 to the turbine case 20.

Turing again to the turbine case 20, the turbine case 20 may include anannular shell 88, a forward bracket 90, and an aft bracket 92 as shownin FIGS. 4-6. The annular shell 88 extends around the axis 11. Theforward bracket 90 extends radially inward from the annular shell 88.The aft bracket 92 extends radially inward from the annular shell 88 ata location axially spaced from the forward bracket 90. The forward andaft brackets 90, 92 also extend circumferentially at least partwayaround the overall circumferential length of the annular shell 88. Inthe illustrative embodiment, the turbine case 20 only has two brackets90, 92. In other embodiments, the turbine case 20 may include two ormore brackets.

In the illustrative embodiment, the forward bracket 90 provides aforward attachment feature for the case mount 77 of the structural vanes36 and the vane mount unit 52 of the composite aero vanes 34 with a hookshape, while the aft bracket 92 provides an aft attachment feature forthe case mount 77 and the vane mount unit 52 with a rail shape. In otherembodiments, the forward and aft brackets 90, 92 may both be hookshaped. In other embodiments, the forward and aft attachment features90, 92 may have another suitable shape (dovetail interface, T-shapeinterface, or other suitable interface shape). Additionally, seals mayalso be arranged between the brackets 90, 92 and the case mount 77and/or the vane mount unit 52 to seal between the components.

In the illustrative embodiment, the forward and aft attachment features90, 92 are axisymmetric about the axis 11. The forward and aftattachment features 90, 92 use the same general attachment and loadtransfer method for both the composite aero and structural vanes 34, 36.

In the illustrative embodiment, the rotatable seal element 32 includes aknife ring 33 as shown in FIG. 4. The knife ring 33 engages the seallands 40, 41 of the static seal element 38 to form the seal 35 betweenthe static seal element 38 and the rotatable seal element 32.

The present disclosure is related to a turbine section 18 of a gasturbine engine 10 including a small number of metallic nozzle guidevanes 36 to carry structural loads that ceramic matrix composite vanes34 cannot tolerate. In the illustrative embodiment, the ceramic matrixcomposite vanes 34 do not carry additional loading. As the ceramicmatrix composite material does not need to carry additional loading, thestresses and/or design flexibility of the composite vanes 34 willimprove.

The turbine vane assembly 26 may be configured to support other gasturbine engine components, such as an inter-stage seal 38. Accordingly,an application of a metallic support structure is likely to be requiredto transmit the axial loading applied to the components to thehigh-pressure turbine casing 20.

The present disclosure incorporates a combination of ceramic matrixcomposite and metallic vanes 34, 36, whereby the inter-stage seal axialload is largely transmitted through the metallic vanes 36. An optionalmetallic spar 62 may be installed inside the ceramic matrix compositevanes 34 and may in some embodiments, be configured to accommodate aportion of the inter-stage seal loading.

The number of metallic vanes e.g. 36 may be minimized as they canrequire greater cooling flows compared to the ceramic matrix compositevanes e.g. 34. However, the number of metallic vanes 36 may be greaterthan 1 to introduce an element of redundancy to the inter-stage sealsupport structure. The metallic vanes 36 may be circumferentiallyequally spaced.

The same aerodynamic definition or airfoil shape could be applied toboth metal and ceramic matrix composite structures. However,consideration of the mismatch in thermal expansion is required,particularly at the ceramic matrix composite-metal vane interface.Achievement of equivalent aerodynamic performance between the twoaerofoil styles could be a consideration to avoid introducing additionalvibration forcing frequencies.

The ceramic matrix composite and metallic vanes 34, 36 may have improvedaerodynamic performance when compared to a uniformly size set of ceramicmatrix composite airfoils. The uniform ceramic matrix composite airfoils34 may have a relatively large maximum thickness to increase and providea sufficient second moment of area. However, a metallic aerofoil e.g. 36can withstand larger mechanical loads so the mixed material set willhave improved aerodynamic freedom i.e. option for reduced thickness andcould result in an aerodynamically superior solution when compared to auniformly size ceramic matrix composite vanes.

The stress of the metal spar (e.g. 62) is proportional to the loading(or number of metallic vanes) divided by the second moment of area ofthe vane, therefore an optimum aerodynamic solution may be to balancethe number of metallic vanes (coolant consumption) against the size ofthe vanes (aerodynamic loss) that just provides an acceptable stress inthe metallic vane.

The damage mechanisms associated with the ceramic matrix compositematerial are less studied than metallics and silicon carbide/siliconcarbide ceramic matrix composites suffer from relatively lowenvironmental durability. This concept takes advantage of thesignificant engine experience associated with metallic nickel vanes(e.g. 36) along with better understanding of the damage mechanisms,providing a robust and reliable support structure. It may beadvantageous to only transmit the flow through the metallic vanes (e.g.36).

While the disclosure has been illustrated and described in detail in theforegoing drawings and description, the same is to be considered asexemplary and not restrictive in character, it being understood thatonly illustrative embodiments thereof have been shown and described andthat all changes and modifications that come within the spirit of thedisclosure are desired to be protected.

What is claimed is:
 1. A turbine section adapted for use in a gasturbine engine, the turbine section comprising a case made from metallicmaterials and shaped to extend around a central reference axis, aturbine wheel housed in the case, the turbine wheel including a diskmounted for rotation about the central reference axis relative to thecase, a plurality of blades coupled to the disk for rotation with thedisk, and a rotatable seal element coupled to the disk for rotation withthe disk, and a turbine vane assembly fixed to the case and configuredto smooth and redirect air moving along a primary gas path of theturbine section ahead of interaction with the turbine wheel, the turbinevane assembly including a plurality of composite aero vanes made fromceramic matrix composite material shaped to provide inner and outer endwalls defining the primary gas path as well as airfoils that extendacross the primary gas path, a plurality of structural vanes made frommetallic materials shaped to provide airfoils that extend across theprimary gas path, and a static seal element that cooperates with therotatable seal element of the turbine wheel to provide a seal forresisting movement of gasses across the seal when the turbine section isin use within a gas turbine engine, wherein the static seal element isfixed to the plurality of structural vanes so as to be in turn coupledto the case while remaining free for relative movement in relation tothe composite aero vanes so that the composite aero vanes aresubstantially free from carrying mechanical loads applied by pressure onthe static seal element to the case.
 2. The turbine section of claim 1,wherein the seal provided by the rotatable seal element and the staticseal element are arranged radially inward of the primary gas path. 3.The turbine section of claim 2, wherein the static seal element includesa seal panel that divides axially adjacent compartments within theturbine section and at least one seal land that is engaged by therotatable seal element.
 4. The turbine section of claim 3, wherein therotatable seal element includes a knife ring that engages the land ofthe static seal element.
 5. The turbine section of claim 1, wherein theplurality of structural vanes each include an inner end wall facing theprimary gas path, an outer end wall spaced radially from the inner endwall and facing the primary gas path, and an airfoil that extends fromthe inner end wall to the outer end wall across the primary gas path. 6.The turbine section of claim 5, wherein the plurality of structuralvanes each include a seal mount that extends radially inwardly from theinner end wall away from the primary gas path and the static sealelement is fixed to the seal mount.
 7. The turbine section of claim 6,wherein the plurality of structural vanes each include a case mount thatextends radially outwardly from the outer end wall away from the primarygas path and the case mount engages the case to couple the structuralvane and the static seal element to the case.
 8. The turbine section ofclaim 7, wherein the plurality of composite aero vanes are each coupledto the case by a metallic spar that extends through the airfoils of thecomposite aero vanes.
 9. The turbine section of claim 6, wherein thestatic seal element is fixed to the seal mount by at least one fastener.10. The turbine section of claim 6, wherein each of the plurality ofcomposite aero vanes is spaced apart from the static seal element.
 11. Aturbine vane assembly adapted for use in a gas turbine engine, theassembly comprising a plurality of composite aero vanes made fromceramic matrix composite material, each of the plurality of compositeaero vanes including an inner end wall that extends partway around acentral reference axis, an outer end wall spaced radially from the innerend wall to define a primary gas path therebetween, and an airfoil thatextends from the inner end wall to the outer end wall across the primarygas path, a plurality of structural vanes made from metallic materials,each of the plurality of structural vanes including airfoils that extendacross the primary gas path, and a static seal element fixed to theplurality of structural vanes and spaced apart from the composite aerovanes.
 12. The assembly of claim 11, wherein the static seal element isarranged radially inward of the primary gas path.
 13. The assembly ofclaim 12, wherein the plurality of structural vanes each include a sealmount that extends radially inwardly from the inner end wall away fromthe primary gas path and the static seal element is fixed to the sealmount.
 14. The assembly of claim 13, wherein the plurality of structuralvanes each include a case mount that extends radially outwardly from theouter end wall away from the primary gas path, and the case mount isconfigured to engage a case so that the static seal element may be fixedto the case via the plurality of structural vanes.
 15. The assembly ofclaim 13, wherein the static seal element is fixed to the seal mount byat least one fastener.
 16. The assembly of claim 13, wherein theplurality of structural vanes each include an inner end wall facing theprimary gas path, an outer end wall spaced radially from the inner endwall and facing the primary gas path, and an airfoil that extends fromthe inner end wall to the outer end wall across the primary gas path.17. The assembly of claim 16, wherein the airfoil of each of theplurality of structural vanes is formed to include a cooling airpassageway and a plurality of film cooling holes interconnecting thecooling air passageway with the primary gas path.
 18. The assembly ofclaim 11, wherein the plurality of composite aero vanes are each mountedto a metallic spar that extends radially through the associated airfoil.